The present invention relates to a combustor assembly of a gas turbine engine. More specifically, the present invention relates to an apparatus and method of cooling a combustor liner of a gas turbine engine.
A gas turbine engine extracts energy from a flow of hot combustion gases. Compressed air is mixed with fuel in a combustor assembly of the gas turbine engine, and the mixture is ignited to produce hot combustion gases. The hot gases flow through the combustor assembly and into a turbine where energy is extracted.
Conventional gas turbine engines use a plurality of combustor assemblies. Each combustor assembly includes a fuel injection system, a combustor liner and a transition duct. Combustion occurs in the combustion liner. Hot combustion gases flow through the combustor liner and the transition duct into the turbine.
The combustor liner, transition duct and other components of the gas turbine engine are subject to these hot combustion gases. Current design criteria require that the combustor liner be cooled to keep its temperature within design parameters. One cooling technique is impingement cooling a surface wall of the combustor liner.
In impingement cooling of a combustor liner, a jet-like flow of cooling air is directed towards the backside wall (outer surface) of the liner, where the front side (inner surface) of the liner is exposed to the hot gases. After impingement, the “spent air” (i.e. air after impingement) flows generally parallel to the combustor liner.
Gas turbine engines may use impingement cooling to cool combustor liners and transition ducts. In such arrangements, the combustor liner is surrounded by a flow sleeve, and the transition duct is surrounded by an impingement sleeve. The flow sleeve and the impingement sleeve are each formed with a plurality of rows of cooling holes.
A first flow annulus is created between the flow sleeve and the combustor liner. The cooling holes in the flow sleeve direct cooling air jets into the first flow annulus to cool the combustor liner. After impingement, the spent air flows axially through the first flow annulus in a direction generally parallel to the combustor liner.
A second flow annulus is created between the transition duct and the impingement sleeve. The holes in the impingement sleeve direct cooling air into the second flow annulus to cool the transition duct. After impingement, the spent air flows axially through the second flow annulus, in a direction generally parallel to the transition duct.
The combustor liner and the transition duct connect, and the flow sleeve and the impingement sleeve connect, such that the first flow annulus and the second flow annulus create a continuous flow path. That is, spent air from the second flow annulus continues into the first flow annulus. This flow from the second flow annulus may reduce the effectiveness and efficiency of the cooling air jets of the flow sleeve. For example, flow through the second flow annulus may bend the jets entering through the flow sleeve, reducing the heat transferring effectiveness of the jets or completely preventing the jets from reaching the surface of the combustor liner. This is especially a problem with regard to the first row of flow sleeve cooling holes adjacent the impingement sleeve.